This invention relates generally to gas turbine engines, and in particular. to a cooled flow path surface region.
This application references co-pending applications assigned to the assignee of the present invention, which are identified as Ser. No. 09/707,023 entitled xe2x80x9cDirectly Cooled Thermal Barrier Coating Systemxe2x80x9d, Ser. No. 09/707,027 entitled xe2x80x9cTranspiration Cooling in Thermal Barrier Coatingxe2x80x9d and Ser. No. 09/707,024 entitled xe2x80x9cMulti-layer Thermal Barrier Coating with Integrated Cooling System,xe2x80x9d the contents of which are incorporated herein by reference.
In gas turbine engines, for example, aircraft engines, air is drawn into the front of the engine, compressed by a shaft-mounted rotary-type compressor, and mixed with fuel. The mixture is burned, and the hot exhaust gases are passed through a turbine mounted on a shaft. The flow of gas turns the turbine, which turns the shaft and drives the compressor and fan. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forward.
During operation of gas turbine engines, the temperatures of combustion gases may exceed 3,000xc2x0 F., considerably higher than the melting temperatures of the metal parts of the engine, which are in contact with these gases. Operation of these engines at gas temperatures that are above the metal part melting temperatures is a well established art, and depends in part on supplying cooling air to the metal parts through various methods. The metal parts of these engines that are particularly subject to high temperatures, and thus require particular attention with respect to cooling, are the metal parts forming combustors and parts located aft of the combustor including turbine blades and vanes and exhaust nozzles.
The hotter the turbine inlet gases, the more efficient is the operation of the jet engine. There is thus an incentive to raise the turbine inlet gas temperature. However, the maximum temperature of the turbine inlet gases is normally limited by the materials used to fabricate the components downstream of the combustors such as the vanes and the blades of the turbine. In current engines, the turbine vanes and blades are made of nickel-based superalloys, and can operate at temperatures of around 2100xc2x0 F.
The metal temperatures can be maintained below melting levels with current cooling techniques by using a combination of improved cooling designs and thermal barrier coatings (TBCs). For example, with regard to the metal blades and vanes employed in aircraft engines, some cooling is achieved through convection by providing passages for flow of cooling air from the compressor internally within the blades so that heat may be removed from the metal structure of the blade by the cooling air. Such blades have intricate serpentine passageways within the structural metal forming the cooling circuits of the blade.
Small internal orifices have also been devised to direct this circulating cooling air directly against certain inner surfaces of the airfoil to obtain cooling of the inner surface by impingement of the cooling air against the surface, a process known as impingement cooling. In addition, an array of small holes extending from a hollow core through the blade shell can provide for bleeding cooling air through the blade shell to the outer surface where a film of such air can protect the blade from direct contact with the hot gases passing through the engine, a process known as film cooling.
In another approach, a thermal barrier coating (TBC) is applied to the turbine blade component, which forms an interface between the metallic component and the hot gases of combustion. The TBC includes a ceramic coating that is applied to the external surface of metal parts to impede the transfer of heat from hot combustion gases to the metal parts, thus insulating the component from the hot combustion gas. This permits the combustion gas to be hotter than would otherwise be possible with the particular material and fabrication process of the component.
TBCs include well-known ceramic materials, for example, yttrium-stabilized zirconia (YSZ). Ceramic TBCs usually do not adhere well directly to the superalloys used as substrate materials. Therefore, an additional metallic layer called a bond coat is placed between the substrate and the TBC. The bond coat may be made of a nickel-containing overlay alloy, such as a MCrAlY, where M is an element selected from the group consisting of Ni, Co, Fe and combinations thereof, or other compositions more resistant to environmental damage than the substrate. Alternatively, the bond coat may be a diffusion nickel aluminide or platinum aluminide, which is grown into the surface of the substrate and whose surface oxidizes to form a protective aluminum oxide scale that provides improved adherence of the ceramic top coatings. The bond coat and overlying TBC are frequently referred to as a thermal barrier coating system.
In an effort to improve TBC life, U.S. Pat. No. 5,419,971 to Skelly et al., assigned to the assignee of the present invention, is directed to small grooves placed in the bond coat layer and/or an interfacial layer lying between the substrate and the TBC to minimize spallation resulting from propagation of cracks formed in TBC systems. The grooves are formed by an ablation process using an ultraviolet laser such as an excimer laser. These grooves have been demonstrated to improve TBC life by facilitating the formation of desired TBC microstructure, which arrests the propagation of cracks formed within TBC, thereby reducing the incidence of spallation, defined as the chipping or flaking away of the coating.
Attempts to improve the life of the bond coat include U.S. Pat. No. 5,034,284 to Bornstein et al. which discloses a porous strain isolation layer placed between the substrate and the bond coat. The porous layer is formed by spraying a mixture of alloy and polymer particles with subsequently heating to eliminate the polymer. The pores are in a random pattern and do not create channels.
The three co-pending applications referenced above disclose small cooling or micro channels within or near the bond coat to improve bond coat and/or TBC system life. These micro channels may communicate directly with at least one cooling circuit contained within the component from which they receive cooling air, thereby providing direct and efficient cooling for the TBC system. To form these micro channels, a surface is masked with a masking material, the masking material forming a pattern on the surface overlying at least one cooling fluid supply contained within the component. The masking material is subsequently removed, leaving hollow micro channels to actively cool the flow path surface region, thus achieving a better, more efficient engine performance.
Creating micro grooves with an excimer laser is a slow and expensive process. Utilizing a masking material which must later be removed also adds additional time and expense. Thus, there is an ongoing need for improved methods for economically creating micro grooves or channels used to encourage favorable microstructure formation and/or improve the environmental resistance and long-term stability of the thermal barrier coating system, so that higher engine efficiencies can be obtained. The present invention fulfills this need, and further provides related advantages.
The present invention provides an improved method for creating micro grooves or channels within or adjacent to the TBC layer applied to a gas turbine engine component, for example, a blade or vane.
In one embodiment, a substrate surface is first coated with a bond coat, for example, an approximately 0.002xe2x80x3 thick diffusion PtAl or an overlay NiAl based alloy coating. A wire mesh is placed a predetermined distance above the bond coat surface, and an inner TBC layer, approximately 0.002xe2x80x3 thick is then deposited on top of the bond coat. The wires in the mesh shadow the TBC deposition, forming structured grooves on the TBC surface. The wire mesh is then removed and an additional, outer, TBC is deposited.
The screen is removed while the shadow-formed grooves are relatively shallow. Subsequent electron beam evaporation physical vapor deposition (EB-PVD) TBC coating is then applied to achieve the microstructure described in the above referenced U.S. Pat. No. 5,419,971, assigned to the assignee of the present invention, incorporated by reference in its entirety herein.
In a different embodiment, the EB-PVD process is used to deposit an outer TBC, for example, a porous, columnar TBC microstructure over an inner TBC, using a screen so as to form channels at the interface between the inner and outer TBC. After grooves are created at the interface, the screen is removed and deposition of the outer TBC is completed, leaving channels in the outer TBC, in a manner similar to that previously set forth above. The composition and microstructure of the outer TBC layer may be different from the inner TBC, and thus controlled as required. The mesh size and wire diameter of the mesh design of the screen may be varied, as can the distance between the bond coat and screen, along with the motion of the screen in the X and Y directions, so as to create shadowed channels/voids of varying geometry.
When placed adjacent to or within a porous TBC, the micro channels provide both active and transpiration cooling through the porous TBC. The micro channels are placed to communicate directly with at least one cooling circuit contained within component from which they receive cooling air, thereby providing direct and efficient cooling for the TBC system. The result is a substrate having an actively cooled flow path surface region that can reduce the cooling requirement for the substrate. Therefore, the engine can run at a higher firing temperature without the need for additional cooling air, achieving a better, more efficient engine performance.
The present invention further comprises the cooled flow path surface region of a component formed by the foregoing methods and the turbine component with the patterned micro channels and/or structured grooves formed by the foregoing methods for cooling the component and/or arresting the propagation of cracks within the TBC.
One advantage of the present invention is that because the wire mesh can be cleaned and reused repeatedly, the wire mesh shadowing technique disclosed is more economical than the laser grooving technique.
Yet another advantage is that the wire size and mesh density can be varied to easily control the size and shape of the structured grooves best suited for various requirements and applications. For example, the structured grooves may be made small enough to act not as cooling channels, but rather as controlled porosity to reduce thermal conductivity, and delay the propagation of TBC spallation; or, alternatively, cooling channels with differing sizes and geometric cross sections may be easily obtained.
Still another advantage is the composition and microstructure of the TBC is easily varied to suit specific requirements and applications. For example, when micro channels are formed (as discussed below), cooling air supplied to these micro channels can provide a direct cooling to the TBC and bond coat to improve the TBC life. By controlling the outer TBC porosity, cooling air will also flow through the porosity inside the ceramic layer to provide transpiration cooling.
Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment, taken in conjunction with the accompanying figures which illustrate, by way of example, the principles of the invention.